Turbofan engine having inner fixed structure including ducted passages

ABSTRACT

A gas turbine engine system includes a fan bypass passage ( 27 ), a core nacelle ( 28 ) having an inner fixed structure ( 40 ) within the fan bypass passage, a passage ( 42 ) extending through the inner fixed structure, and a duct nozzle ( 48 ). The passage includes an inlet ( 44 ) for receiving a fan airflow (F 2 ) from the fan bypass passage and an outlet ( 46 ) for discharging the fan airflow. The duct nozzle includes a variable cross-sectional exit area ( 50 ) for controlling the fan airflow within the passage and is selectively moveable to influence the variable cross-sectional exit area.

BACKGROUND OF THE INVENTION

This invention generally relates to a gas turbine engine, and moreparticularly to a turbofan gas turbine engine having a ducted innerfixed structure for modulating a discharge airflow cross-sectional areaof the gas turbine engine.

In an aircraft gas turbine engine, such as a turbofan engine, air ispressurized in a compressor, and mixed with fuel and burned in acombustor for generating hot combustion gases. The hot combustion gasesflow downstream through turbine stages that extract energy from thegases. A high pressure turbine powers the compressor, while a lowpressure turbine powers a fan section disposed upstream of thecompressor.

Combustion gases are discharged from the turbofan engine through a coreexhaust nozzle, and fan air is discharged through an annular fan exhaustnozzle defined at least partially by a fan nacelle surrounding the coreengine. A significant amount of propulsion thrust is provided by thepressurized fan air which is discharged through the fan exhaust nozzle.The combustion gases are discharged through the core exhaust nozzle toprovide additional thrust.

A significant amount of the air pressurized by the fan section bypassesthe engine for generating propulsion thrust in turbofan engines. Highbypass turbofans typically require large diameter fans to achieveadequate turbofan engine efficiency. Therefore, the nacelle of theturbofan engine must be large enough to support the large diameter fanof the turbofan engine. Disadvantageously, the relatively large size ofthe nacelle results in increased weight, noise and drag that may offsetthe propulsive efficiency achieved by the high bypass turbofan engine.

It is known in the field of aircraft gas turbine engines that theperformance of the turbofan engine varies during diverse flightconditions experienced by the aircraft. Typical turbofan engines aredesigned to achieve maximum performance during normal cruise operationof the aircraft. Therefore, when combined with the necessity of arelatively large nacelle size, increased noise and decreased efficiencymay be experienced by the aircraft at non-cruise operating conditionssuch as take-off, landing, cruise maneuver and the like.

Accordingly, it is desirable to provide a turbofan engine having adischarge airflow cross-sectional area that may be modulated to achievenoise reductions and improved safety and efficiency of the gas turbineengine in a relatively inexpensive and non-complex manner.

SUMMARY OF THE INVENTION

An example gas turbine engine system includes a fan bypass passage, acore nacelle having an inner fixed structure within the fan bypasspassage, a passage extending through the inner fixed structure, and aduct nozzle. The passage includes an inlet for receiving a fan airflowfrom the fan bypass passage and an outlet for discharging the fanairflow. The duct nozzle includes a variable cross-sectional exit areafor controlling the fan airflow within the passage and is selectivelymoveable to influence the variable cross-sectional exit area.

A second example gas turbine engine system includes a fan nacelle, acore nacelle having an inner fixed structure, a fan bypass passage, apassage extending through the inner fixed structure, a duct nozzle, afan section, at least one compressor and at least one turbine, acombustor, a sensor that produces a signal representing an operabilitycondition and a controller that receives the signal. The passageincludes an inlet for receiving a fan airflow from the fan bypasspassage and an outlet for discharging the fan airflow. The duct nozzleis selectively moveable to vary a variable cross-sectional exit area inresponse to an operability condition.

An example method of modulating a variable cross-sectional exit area ofa gas turbine engine system includes sensing an operability conditionand selectively varying a cross-sectional exist area of a duct nozzle ofa passage in response to sensing the operability condition.

The various features and advantages of this invention will becomeapparent to those skilled in the art from the following detaileddescription. The drawings that accompany the detailed description can bebriefly described as follows.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 illustrates a general prospective view of an example gas turbineengine;

FIG. 2 is an exploded schematic view of an example gas turbine enginehaving an inner fixed structure including a ducted passage having avariable cross-sectional exit area;

FIG. 3 illustrates an example gas turbine engine having a plurality ofducted passages;

FIG. 4 is an end view of a second example gas turbine engine having aplurality of ducted passages;

FIG. 5 is an end view of another example gas turbine engine having aplurality of ducted passages;

FIG. 6 is an end view of yet another example gas turbine engine having aplurality of ducted passages;

FIG. 7 illustrates an example configuration of an outlet of the ductedpassage of the gas turbine engine;

FIG. 8 illustrates an example configuration of a duct nozzle associatedwith the ducted passage of the gas turbine engine; and

FIG. 9 is a schematic view of an inlet of the ducted passage of the gasturbine engine.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT

Referring to FIG. 1, a gas turbine engine 10 suspends from an enginepylon 12 as is typical of an aircraft designed for subsonic operation.In one example, the gas turbine engine is a geared turbofan aircraftengine. The gas turbine engine 10 includes a fan section 14, a lowpressure compressor 15, a high pressure compressor 16, a combustor 18, ahigh pressure turbine 20 and a low pressure turbine 22. A low speedshaft 19 rotationally supports the low pressure compressor 15 and thelow pressure turbine 22 and drives the fan section 14 through a geartrain 23. A high speed shaft 21 rotationally supports the high pressurecompressor 16 and a high pressure turbine 20. The low speed shaft 19 andthe high speed shaft 21 rotate about a longitudinal centerline axis A ofthe gas turbine engine 10.

During operation, air is pressurized in the compressors 15, 16 and ismixed with fuel and burned in the combustor 18 for generating hotcombustion gases. The hot combustion gases flow through the high and lowpressure turbines 20, 22 which extract energy from the hot combustiongases.

The example gas turbine engine 10 is in the form of a high bypass ratio(i.e., low fan pressure ratio geared) turbofan engine mounted within afan nacelle 26, in which most of the air pressurized by the fan section14 bypasses the core engine itself for the generation of propulsionthrust. The example illustrated in FIG. 1 depicts a high bypass flowarrangement in which approximately 80% of the airflow entering the fannacelle 26 may bypass the core nacelle 28 via a fan bypass passage 27.The high bypass flow arrangement provides a significant amount of thrustfor powering the aircraft.

In one example, the bypass ratio is greater than ten to one, and the fansection 14 diameter is substantially larger than the diameter of the lowpressure compressor 15. The low pressure turbine 22 has a pressure ratiothat is greater than five to one, in one example. The gear train 23 canbe any known gear system, such as a planetary gear system with orbitingplanet gears, planetary system with non-orbiting planet gears, or othertype of gear system. In the disclosed example, the gear train 23 has aconstant gear ratio. It should be understood, however, that the aboveparameters are only exemplary of a contemplated geared turbofan engine.That is, the invention is applicable to other engine architectures,including direct drive turbofans.

A fan discharge airflow F1 is communicated within the fan bypass passage27 and is discharged from the engine 10 through a fan exhaust nozzle 30,defined radially between a core nacelle 28 and the fan nacelle 26. Coreexhaust gases C are discharged from the core nacelle 28 through a coreexhaust nozzle 32 defined between the core nacelle 28 and a tail cone 34disposed coaxially therein around the longitudinal centerline axis A ofthe gas turbine engine 10.

The fan exhaust nozzle 30 concentrically surrounds the core nacelle 28near an aftmost segment 29 of the fan nacelle 26, in this example. Inother examples, the fan exhaust nozzle 30 is located farther upstreambut aft of the fan section 14. The fan exhaust nozzle 30 defines adischarge airflow cross-sectional area 36 between the fan nacelle 26 andthe core nacelle 28 for axially discharging the fan discharge airflow F1pressurized by the upstream fan section 14.

The core nacelle 28 of the gas turbine engine 10 includes a core cowl38. The core cowl 38 represents an exterior flow surface of a section ofthe core nacelle 28. The core cowl 38 is positioned adjacent an innerduct boundary 25 of the fan bypass passage 27.

FIG. 2 illustrates an inner fixed structure (IFS) 40 of the core nacelle28. The IFS 40 represents a section of hardware of the core nacelle 28that includes the entire inner-duct boundary 25 and the core cowl 38 ofthe core nacelle 28, in one example. The example IFS 40 includes aducted passage 42 which extends therethrough. In one example, the IFS 40includes a plurality of ducted passages 42 disposed circumferentiallyabout the engine centerline axis A (see FIG. 3).

In one example, the ducted passages 42 have a generally oval shapedcross-section (see FIG. 4). In another example, the ducted passages 42have a generally circular shaped cross-section (see FIG. 5). In yetanother example, the duct passages 42 have a generally crescent shapedcross-section (see FIG. 6). The actual number and shape of the ductedpassages 42 will depend upon design specific parameters including, butnot limited to, the size of the core nacelle 28, the arrangement ofengine components located within the IFS 40 and the efficiencyrequirements of the engine 10.

Each ducted passage 42 includes an inlet 44 and an outlet 46. The inlet44 is positioned adjacent a forward section 68 of the fan bypass passage27 (See FIG. 9). The outlet 46 of each ducted passages 42 is positionedjust aft (i.e., downstream) of the fan nacelle 26, in one example (SeeFIG. 7). The actual positioning and configuration of the inlets 44 andthe outlets 46 of the ducted passages 42 will vary depending upon designspecific parameters including, but not limited to, the size of the corenacelle 28, the arrangement of engine components located within the IFS40 and the efficiency requirements of the gas turbine engine 10.

Each inlet 44 receives a portion F2 of the fan airflow F1 from the fanbypass passage 27 as the fan airflow F1 is communicated through the fanbypass passage 27. The airflow F2 is communicated through the ductedpassage 42 and is discharged via the outlet 46.

In the illustrated example, the discharge airflow cross-sectional area36 of the engine 10 extends between the aftmost segment 29 of the fannacelle 26 (i.e., adjacent to the fan exhaust nozzle 30) and the corecowl 38. Modulating the discharge airflow cross-sectional area 36 of thegas turbine engine 10 during specific flight conditions providesadditional fan airflow F1 through the fan bypass passage 27. Dependingupon the application, the additional fan airflow F1 may improveefficiency and reduce noise associated with the gas turbine engine 10.The ducted passages 42 each include a duct nozzle 48 positioned at eachoutlet 46 of the ducted passages 42. In one example, the duct nozzle 48includes a flap 62. Each flap 62 of the duct nozzle 48 is selectivelymoved to control the airflow F2 within each ducted passage 42. In oneexample, a cross-sectional exit area 50 of the outlet 46 of each ductedpassage 42 is varied to provide additional area to the discharge airflowcross-sectional area 36.

The flaps 62 of each duct nozzle 48 are selectively actuated to controlthe air pressure of the fan airflow F1 within the fan bypass passage 27.For example, closing the flaps 62 reduces the cross-sectional exit area50, which restricts the fan airflow F1 and produces a pressure build-up(i.e., an increase in air pressure) within the fan bypass passage 27.Opening the flaps 62 increases the cross-sectional exit area 50, whichpermits more fan airflow F1 and reduces the pressure build-up (i.e., adecrease in air pressure) within the fan bypass passage 27.

The flap 62 of each duct nozzle 48 is moved from a first position X(i.e., a closed position, represented by solid lines) to a secondposition X′ (an open position, represented by phantom lines) in responseto detecting an operability condition of the gas turbine engine 10, inone example. In another example, the flap 62 is selectively moveablebetween a plurality of positions each having different cross-sectionalexit areas associated therewith.

In the illustrated example, the cross-sectional exit area 50 of the ductnozzle 48 associated with the second position X′ is greater than thecross-sectional exit area of the duct nozzle 48 associated with thefirst position X. Movement of the duct nozzle 48 to the second positionX′ to provide the cross-sectional exit area 50, in combination with thedischarge airflow cross-sectional area 36 associated with the fan bypasspassage 27, permits an increased amount of fan airflow F1 to exit thegas turbine engine 10 as compared to the discharge airflow crosssectional area 36 alone. Therefore, the design of the fan section 14 maybe optimized for diverse operability conditions of the aircraft.

In one example, the operability condition includes a low powercondition. Low power conditions include idle conditions, fly-idleconditions and approach conditions, such as where the aircraft isdescending to prepare to land. In another example, the operabilitycondition includes static conditions. Static conditions include takeoffconditions and any other ground operations of the aircraft. However, theduct nozzle 48 may be moved between the first position X and the secondposition X′, or to any other position between the first position X andthe second position X′, in response to any known operability condition.

A sensor 52 detects the operability condition and communicates with acontroller 54 to move the flap 62 of the duct nozzle 48 between thefirst position X and the second position X′ via an actuator assembly 56.Of course, this view is highly schematic. It should be understood thatthe sensor 52 and the controller 54 are programmable to detect knownflight conditions and/or gas turbine engine operability conditions. Aperson of ordinary skill in the art having the benefit of the teachingsherein would be able to program the controller 54 to communicate withthe actuator assembly 56 to move the flap 62 of the duct nozzle 48between the first position X and the second position X′. The actuatorassembly 56 returns the flap 62 to the first position X during normalcruise operation (e.g., a generally constant speed at a generallyconstant, elevated altitude).

FIG. 8 illustrates an example configuration of the duct nozzle 48 ofeach ducted passage 42. In one example, the duct nozzle 48 generallyincludes a synchronizing ring 58, a static ring 60 and at least one flap62. The flap 62 is pivotally mounted to the static ring 60 via a hinge64 and linked to the synchronizing ring 58 through a linkage 66. In theillustrated example, the actuator 56, the synchronizing ring 58, thestatic ring 60, the hinges 64 and the linkage 66 are each enclosedwithin the IFS 40 (i.e., outside of the flowpath of fan airflow F1). Inanother example, the flaps 62 are flush with the core cowl 38 when in aclosed position.

The actuator assembly 56 selectively rotates the synchronizing ring 58relative to the static ring 60 to adjust the flap 62 through the linkage66. The radial movement of the synchronizing ring 58 is converted totangential movement of the flap 62 to vary the cross-sectional exit area50 of the duct nozzle 48 and permit/restrict the discharge of theairflow F2 via the outlet 46.

FIG. 9 illustrates an example configuration of the inlet 44 of eachducted passage 42. In one example, the inlet 44 of each ducted passage42 is positioned adjacent a forward section 68 of the fan bypass passage27. It should be understood that the inlet 44 of the ducted passages 42may be positioned at any other location relative to the fan bypasspassage 27.

In one example, each inlet 44 is designed with an inlet shape thateliminates spillage when the duct nozzle 48 is in a fully closedposition (i.e., the first position X). Airflow spillage occurs inresponse to the fan airflow F1 partially entering the inlet 44 andsubsequently “spilling” out of the inlet 44 and back into the fan bypasspassage 27 where the duct nozzle 48 is in the fully closed position X(See FIG. 2). Airflow spillage creates the potential for system thrustloss caused by the turbulence of the fan airflow that “spills” out ofthe inlet and back into the fan bypass passage 27.

During different flight conditions, the amount of fan airflow F1 whichis communicated through the fan bypass passage 27 varies. Therefore, theinlet 44 of the ducted passage 42 must be suitably shaped to reduceairflow spillage as the fan airflow is communicated through the fanbypass passage 27 and the duct nozzle 48 is closed.

In another example, the inlet 44 optionally includes an inlet door 70that reduces the risk of airflow spillage. In the illustrated example,the inlet door 70 is fully closed when the duct nozzle 48 is fullyclosed, and fully open when the duct nozzle 48 is in any open position,such as the second position X′, for example. That is, where the ductnozzle 48 is positioned at the first position X, the inlet door 70 isclosed to prevent any fan airflow F1 from being communicated from thefan bypass passage 27 into the ducted passage 42.

The inlet door 70 is mounted at its leading edge 72 to the core nacelle28. In one example, the inlet door is pivotable about the leading edge72 to open and close the inlet 44 of the ducted passage 42. A worker ofordinary skill in the art having the benefit of this disclosure would beable to pivotally mount the inlet door 70 to the inlet 44 to reduce therisk of airflow spillage.

The foregoing description shall be interpreted as illustrative and notin any limiting sense. A worker of ordinary skill in the art wouldrecognize that certain modifications would come within the scope of thisinvention. For that reason, the follow claims should be studied todetermine the true scope and content of this invention.

1. A gas turbine engine system, comprising: a fan bypass passage; a corenacelle having an inner fixed structure within the fan bypass passage; aplurality of passages extending through said inner fixed structure andbeing circumferentially spaced about an engine centerline axis of saidgas turbine engine, wherein each of said plurality of passages includesan inlet for receiving a fan airflow from said fan bypass passage and anoutlet for discharging said fan airflow, and individual duct walls thatextend between said inlet and said outlet of each of said plurality ofpassages, wherein each of said individual duct walls of said pluralityof passages are circumferentially spaced apart; and a duct nozzle havinga variable cross-sectional exit area for controlling said fan airflowwithin each of said plurality of passages, wherein said duct nozzle isselectively moveable to influence said variable cross-sectional exitarea.
 2. The system as recited in claim 1, wherein said duct nozzle ismounted to said outlet of each of said plurality of passages, said ductnozzle comprising at least one flap that pivots about a hinge.
 3. Thesystem as recited in claim 1, wherein said inlet of each of saidplurality of passages is positioned adjacent a forward section of saidfan bypass passage, said inlet having an access door to reduce spillageof said fan airflow as said fan airflow is communicated through said fanbypass passage.
 4. The system as recited in claim 1, wherein said outletof each of said plurality of passages is positioned downstream from afan nacelle which at least partially surrounds said core nacelle.
 5. Thesystem as recited in claim 1, wherein each of said plurality of passagesincludes at least one of a generally oval shaped cross-section, agenerally circular shaped cross-section, and a generally crescent shapedcross-section.
 6. The system as recited in claim 1, wherein said ductnozzle includes a synchronizing ring, a static ring, at least one flappivotable about a hinge and a linkage, wherein an actuator assemblyselectively rotates said synchronizing ring relative to said static ringto adjust said at least one flap through said linkage.
 7. A gas turbineengine system, comprising: a fan nacelle defined about an axis andhaving a fan exhaust nozzle; a core nacelle having an inner fixedstructure at least partially within said fan nacelle; a fan bypasspassage between said fan nacelle and said core nacelle; a plurality ofpassages extending through said inner fixed structure andcircumferentially disposed about said axis, wherein each of saidplurality of passages includes an inlet for receiving a fan airflow fromsaid fan bypass passage and an outlet for discharging said fan airflow,and individual duct walls that extend between said inlet and said outletof each of said plurality of passages, wherein each of said individualduct walls of said plurality of passages are circumferentially spacedapart; a duct nozzle having a variable cross-sectional exit area forcontrolling said fan airflow within each of said plurality of passages,wherein said duct nozzle is selectively moveable to vary said variablecross-sectional exit area in response to an operability condition; a fansection positioned within said fan nacelle; at least one compressor andat least one turbine positioned downstream of said fan section; at leastone combustor positioned between said at least one compressor and saidat least one turbine; at least one sensor that produces a signalrepresenting said operability condition; and a controller that receivessaid signal, wherein said controller selectively moves said duct nozzlein response to said signal.
 8. The gas turbine engine system as recitedin claim 7, wherein said duct nozzle is mounted to said outlet of eachof said plurality of passages downstream from said fan nacelle.
 9. Thegas turbine engine system as recited in claim 7, wherein said ductnozzle is moveable between a first position having a firstcross-sectional exit area and a second position having a secondcross-sectional exit area greater than said first cross-sectional exitarea in response to said signal.
 10. The gas turbine engine system asrecited in claim 7, wherein said operability condition includes at leastone of a low power condition and a static condition.
 11. The gas turbineengine system as recited in claim 7, comprising an actuator assembly incommunication with said controller for selectively moving said ductnozzle.
 12. The gas turbine engine system as recited in claim 11,wherein said duct nozzle includes a synchronizing ring, a static ring,at least one flap pivotable about a hinge and a linkage, wherein saidactuator assembly selectively rotates said synchronizing ring relativeto said static ring to adjust said at least one flap through saidlinkage.
 13. A method of modulating a discharge-airflow cross-sectionalarea of a gas turbine engine system including a fan bypass passage andan inner fixed structure including a plurality of passagescircumferentially disposed about an engine centerline axis, each havingan inlet for receiving a fan airflow from the fan bypass passage and anoutlet for discharging the fan airflow, and individual duct walls thatextend between said inlet and said outlet of each of said plurality ofpassages, wherein each of said individual duct walls of said pluralityof passages are circumferentially spaced apart, comprising the steps of:(a) sensing an operability condition; and (b) selectively varying across-sectional exit area of a duct nozzle of each of the plurality ofpassages in response to sensing the operability condition.
 14. Themethod as recited in claim 13, wherein the duct nozzle is moveablebetween a first position having a first cross-sectional exit area and asecond position having a second cross-sectional exit area greater thanthe first cross-sectional exit area, wherein said step (b) furthercomprises: moving the duct nozzle between the first position and thesecond position.
 15. The method as recited in claim 14, comprising thestep of: (c) returning the duct nozzle to the first position in responseto detection of a cruise operation.
 16. The method as recited in claim13, wherein the operability condition includes at least one of a lowpower condition and a static condition.
 17. The method as recited inclaim 13, wherein said step (b) further comprises: mounting at least oneflap to the outlet of each of the plurality of passages; and selectivelypivoting the at least one flap to vary the cross-sectional exit area ofthe duct nozzle.